A compressor or turbine rotor assembly for gas turbine engines, includes a plurality of compressor or turbine blades mounted to the outer periphery of a central rotor. It is important in gas turbine engines that the outer tips of the compressor or turbine blades run very close to surrounding shrouds in order to minimize gas leakage across the tips of the blades. Machining of such compressor or turbine blade tips to the desired outer true tip diameter, is a difficult manufacturing operation because the blades are normally retained with root sections loosely fitting within dovetail grooves in the rotor at the periphery thereof. Prior art fixture tools use radial expansion of resilient materials under axial compression forces for simulating the centrifugal force created during engine operation, to radially position the blades during a machining process. In such a prior art method, it is difficult to accurately control the quantity, acting points, and even distribution of radial forces acting on the individual blades. Therefore, the results are often unsatisfactory. In another prior art machining method, the rotor assembly is rotated at a high speed, resulting in a centrifugal force for positioning the blades within the slots of the rotor in a blade tip machining process. The high speed rotation of the rotor disc during the machining process is however, not desirable due to various concerns such as safety, convenience, cost of the manufacturing process, etc.
Accordingly, there is a need to provide an improved apparatus and method for machining blade tips as used in gas turbine engines.